(1) Field of the Invention
The present invention relates to a turbine engine component having an airfoil portion with a serpentine cooling microcircuit embedded in the pressure side, which serpentine cooling microcircuit is provided with a way to increase coolant pressure and a way to accelerate local cooling flow and increase the ability to pick-up heat.
(2) Prior Art
The overall cooling effectiveness is a measure used to determine the cooling characteristics of a particular design. The ideal non-achievable goal is unity, which implies that the metal temperature is the same as the coolant temperature inside an airfoil. The opposite can also occur when the cooling effectiveness is zero implying that the metal temperature is the same as the gas temperature. In that case, the blade material will certainly melt and burn away. In general, existing cooling technology allows the cooling effectiveness to be between 0.5 and 0.6. More advanced technology such as supercooling should be between 0.6 and 0.7. Microcircuit cooling as the most advanced cooling technology in existence today can be made to produce cooling effectiveness higher than 0.7.
FIG. 1 shows a durability map of cooling effectiveness (x-axis) vs. the film effectiveness (y-axis) for different lines of convective efficiency. Placed in the map is a point 10 related to a new advanced serpentine microcircuit shown in FIGS. 2a-2c. This serpentine microcircuit includes a pressure side serpentine circuit 20 and a suction side serpentine circuit 22 embedded in the airfoil walls 24 and 26.
The Table I below provides the operational parameters used to plot the design point in the durability map.
TABLE IOperational Parameters forserpentine microcircuitBeta2.898Tg2581 [F]Tc1365 [F]Tm2050 [F]Tm_bulk1709 [F]Phi_loc0.437Phi_bulk0.717Tco1640 [F]Tci1090 [F]eta_c_loc0.573eta_f0.296Total Cooling3.503%Flow10.8WAELegend for Table IBeta = heat loadPhi_loc = local cooling effectivenessPhi_bulk = bulk cooling effectivenessEta_c_loc = local cooling efficiencyEta_f = film effectivenessTg = gas temperatureTc = coolant temperatureTm = metal temperatureTm_bulk = bulk metal temperatureTco = exit coolant temperatureTci = inlet coolant temperatureWAE = compressor engine flow, pps
It should be noted that the overall cooling effectiveness from the table is 0.717 for a film effectiveness of 0.296 and a convective efficiency (or ability to pick-up heat) of 0.573. Also note that the corresponding cooling flow for a turbine blade having this cooling microcircuit is 3.5% engine flow. FIG. 3 illustrates the cooling flow distribution for a turbine blade with the serpentine microcircuits of FIGS. 2a-2c embedded in the airfoils walls.
It should be noted from FIG. 3 that the flow passing through the pressure side serpentine microcircuit 20 is 1.165% WAE (compressor engine flow) in comparison with 0.428 WAE for the suction side serpentine microcircuit 22. This represents a 2.7 fold increase in cooling flow relative to the suction side microcircuit. The reason for this increase stems from the fact that the thermal load to the part is considerably higher for the airfoil pressure side. As a result, the height of the microcircuit channel should be 1.8 fold increase over that of the suction side. That is 0.022 inches vs. 0.012 inches. Besides the increased flow requirement, the driving potential in terms of source to sink pressures for the pressure side circuit 20 is not as high as that for the suction side circuit 22. In considering the coolant pressure on the pressure side circuit 20, at the end of the third or outlet leg, the back flow margin, as a measure of internal to external pressure, is low. As a consequence of this back flow issue, the metal temperature increases beyond the required metal temperature close to the third leg of the pressure side circuit 20. It is desirable to eliminate this problem.